Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics

ABSTRACT

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot. 
     In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

RELATED APPLICATION

This application is a divisional application and claims the benefit ofU.S. Non-provisional application Ser. No. 11/124,592, filed May 6, 2005under 35 U.S.C. 121.

ORIGIN OF THE INVENTION

This invention was made by employees of the United States Government andcontractors operating under a contract with the United StatesGovernment, and thus may be used by or for the Government forgovernmental purposes without the payment of any royalties thereon ortherefor.

INCORPORATION BY REFERENCE

This application incorporates by reference U.S. Non-provisionalapplication Ser. No. 11/124,592 filed May 6, 2005.

BACKGROUND OF THE INVENTION

1. Field of Invention

This invention relates generally to the field of man-made earthsatellites and more specifically to a novel method and associatedapparatus for capturing, servicing and de-orbiting such satellites.

2. Background of the Invention

The last half-century has been a watershed for the development andimplementation of earth-orbit satellites for various purposes. Differenttypes of earth satellites include those designed for communications,earth remote sensing, weather, global positioning, and scientificresearch. A typical example of a communications satellite might beEchoStar 3, which is used to send television signals to homes throughoutNorth America. Communication satellites act as relay stations in space.They are used to “bounce” digital messages, such as telephone calls,television pictures and internet connections, from one part of the worldto another. EchoStar 3 and many other communications satellites are ingeosynchronous orbit. There are more than 100 such communicationsatellites currently orbiting earth.

Earth remote sensing satellites, such as the LandSat series (LandSat 1through LandSat 7), study the earth's surface. From 300 miles above theearth's surface, and more, these satellites use powerful cameras to scanthe planet. Their instruments study earth's plant cover, chemicalcomposition, surface temperature, ice cap thicknesses, and many otherearth systems and features. Such data is useful in vital industries suchas farming, fishing, mining, and forestry. Moreover, remote sensingsatellites are used to study changes in global environments caused byman. Examples of this include areas that are turning into desert(desertification), and the destruction of rain forests (deforestation).

TIROS (Television Infrared Observational Satellite), operated by NOAA(National Oceanic and Atmospheric Administration), is a representativeweather satellite. TIROS is one of several weather satellites making upa system operated by NOAA, which produces data used to forecast weather,track storms, and otherwise engage in meteorological research. There aretwo TIROS satellites circling earth over the poles while satellites fromanother part of the system, the Geostationary Operational EnvironmentalSatellites (GOES), operate in geosynchronous orbit. Using this system ofsatellites, meteorologists study the weather and climate patterns aroundthe world, such as temperature, moisture, and solar radiation in theatmosphere. Also, many weather satellites are equipped with sensors thataid search and rescue operations.

The Global Positioning System (GPS) satellites are in geosynchronousorbit and are able to identify latitude, longitude and altitude withgreat precision. Originally developed by the military, GPS satellitesare now used by a wide variety of people everywhere to find their exactposition. Airplanes, boats, cars and virtually any transportationapparatus are equipped with GPS receivers. Even hand-held GPS deviceshave become a norm with pedestrians and other travelers.

While there are many examples of satellites engaged in scientificresearch, Hubble Space Telescope (HST) is perhaps the best known. Since1990, when HST was launched, the world has had access to remarkablevisual images that have significantly advanced our understanding of thecosmos. HST's optics, science instruments and spacecraft systems worktogether to capture light from far reaches of the universe, convert itinto digital data, and transmit it back to earth. Because application ofthe current invention to HST represents an exemplary mode contemplatedof carrying out the invention at the time of filing for a United Statespatent, an overview of HST's systems is appropriate. However, oneskilled in the art will recognize that the method of this invention canbe applied and adapted to virtually any man made earth satellite.

HST optics are collectively called the Optical Telescope Assembly (OTA),which includes two mirrors, support trusses and the apertures of theaccompanying instruments. OTA's configuration is that of a well-known,straightforward design known as Ritchie-Chretien Cassegrain, in whichthe two specialized mirrors form focused images over the largestpossible field of view. Referring now to FIGS. 1 through 5, HST and OTAare graphically illustrated. While FIG. 1 shows a three dimensionalcutaway view of HST, FIGS. 2 and 3 specifically illustrate OTA and theprinciples of its operation. As best seen in FIG. 3, light entersthrough the main aperture and travels down a tube fitted with bafflesthat keep out stray light. The light is collected by the concave primarymirror and reflected toward the smaller, convex secondary mirror. Thesecondary mirror bounces the light back toward the primary mirror andthrough a smaller aperture in its center. The light is then focused on asmall area called the focal plane, where it is detected by the variousscience instruments.

OTA's mirrors are necessarily very smooth and have precisely shapedreflecting surfaces. They were ground so that their surfaces do notdeviate from a mathematically perfect curve by more than 1/800,000 of aninch. According to this precision, if the primary mirror were scaled upto the diameter of the earth, the biggest bump would be only six inchestall. Shortly after HST was deployed, it was discovered that the curveto which primary mirror was ground was incorrect, causing a sphericalaberration. Fortunately, corrective optics, much like a contact lens,were able to solve this problem.

The mirrors are made of ultra-low expansion glass and kept at a nearlyconstant room temperature (about 70 degrees Fahrenheit) to avoidwarping. The reflecting surfaces are coated with a 3/1,000,000-inchlayer of pure aluminum and protected by a 1/1,000,000-inch layer ofmagnesium fluoride, which makes the mirrors more reflective ofultraviolet light.

HST contains five science instruments, namely the Advanced Camera forSurveys (ACS), the Wide Field and Planetary Camera (WFPC2), the NearInfrared Camera and Multi-Object Spectrometer (NICMOS), the SpaceTelescope Imaging Spectrograph (STIS), and the Fine Guidance Sensors(FGS). These instruments work either together or individually to observethe universe in a unique way.

Installed in the latest space shuttle servicing mission in March of2002, the ACS represents the third generation of science instrumentsflown aboard HST. It occupies the space vacated by the Faint ObjectCamera, HST's “zoom lens” for nearly twelve years. Among other tasks,ACS is used to map distribution of dark matter, detect the most distantobjects in the universe, search for massive planets in other solarsystems, observe weather on other planets in our solar system, and studythe nature and distribution of galaxies. With its wider field of view,sharper image quality, and enhanced sensitivity, ACS expands HST'scapabilities significantly; its cutting edge technology makes HST tentimes more effective and prolongs its useful life. Designed to studysome of the earliest activity in the universe, ACS detectselectromagnetic waves in wavelengths ranging from far ultraviolet toinfrared.

On the inside, ACS is actually a team of three different cameras,specifically the wide field camera, the high-resolution camera, and thesolar blind camera. The wide field camera conducts broad surveys of theuniverse, which reveal clues about how our universe evolved. Incontrast, the high-resolution camera takes detailed pictures of theinner regions of galaxies. The solar blind camera, which blocks visiblelight to enhance ultraviolet sensitivity, focuses on hot stars radiatingin ultraviolet wavelengths.

HST's “workhorse” instrument, WFPC2, is behind most of the famous imagesit produces. This main camera includes 48 filters mounted on four filterwheels, which allow scientists to study precise wavelengths of light andto sense a range of wavelengths from ultraviolet to near-infrared light.Four postage stamp-sized pieces of circuitry called Charge-CoupledDevices (CCDs) collect and record information from stars and galaxies tomake photographs. These detectors are very sensitive to the extremelyfaint light of distant galaxies. In fact, CCDs can see objects that areone billion times fainter than the human eye can see. Less sensitiveCCDs are now found in some videocassette recorders and virtually all newdigital cameras. Each of the four CCDs on HST contains 640,000 pixels.The light collected by each pixel is translated into a number. Thesenumbers are then transmitted to ground-based computers, which convertthem into an image.

NICMOS is HST's “heat sensor” and can see objects in deepestspace—objects whose light takes billions of years to reach earth. Theinstrument's three cameras, each with different fields of view, aredesigned to detect near-infrared wavelengths, which are slightly longerthan the wavelengths of visible light. Much information about the birthof stars, solar systems, and galaxies are revealed in infrared light,which can penetrate the interstellar dust and gas that often blockvisible light. In addition, light from the most distant objects in theuniverse “shifts” into the infrared wavelengths, and so by studyingobjects and phenomena in this spectral region, astronomers can probe thepast, learning how galaxies, stars and planetary systems form.

Just as a camera for recording visible light must be dark inside toavoid exposure to unwanted light, so must a camera for recordinginfrared light be cold inside to avoid unwanted exposure to unwantedlight in the form of heat. To ensure that NICMOS is recording infraredlight from space rather than heat created by its own electronics, itssensitive infrared sensors must operate at very cold temperatures—below77 degrees Kelvin (−321 degrees Fahrenheit). The instrument's detectorswere initially cooled inside a cryogenic dewar (a thermally insulatedcontainer much like a thermos bottle), which contained a 230 pound blockof nitrogen ice. While successful for about two years, the nitrogen icecube melted prematurely. NICMOS was re-chilled during the last HSTservicing mission of March 2002 with a “cryocooler,” an apparatus thatoperates much like a household refrigerator.

STIS in essence acts like a prism to separate detected light into itscomponent colors. This spectrograph instrument thus provides a“fingerprint” of the object being observed, which reveals informationabout its temperature, chemical composition, density and motion.Spectrographic observations also show changes in celestial objects asthe universe evolves. STIS spans ultraviolet, visible and near-infraredwavelengths. Among other tasks, STIS is used to search for black holes.The light emitted by stars and gas orbiting the center of a galaxyappears redder when moving away from earth (redshift) and bluer whencoming toward earth (blueshift). Thus, STIS looks for redshiftedmaterial on one side of the suspected black hole and blueshiftedmaterial on the other, indicating that this material is orbiting at avery high rate of speed, as would be expected when a black hole ispresent. STIS can sample 500 points along a celestial objectsimultaneously, meaning that many regions in planet's atmosphere or manystars within a galaxy can be recorded in one exposure. STIS wasinstalled on HST during the 1997 shuttle servicing mission.

HST's Fine Guidance Sensors, its targeting cameras, provide feedbackused to maneuver the telescope and perform celestial measurements. Whiletwo of the sensors point the telescope at a desired astronomical target,and then hold that target in an instrument's field of view, the thirdsensor is able to perform scientific observations. The FGS aim HST bylocking onto “guide stars” and continuously measuring the position ofthe telescope relative to the object being viewed. Adjustments based onthese constant, minute measurements keep HST pointed in the desireddirection with an accuracy of 0.01 arcsec. The FGS detect when HSTdrifts even the smallest amount and return it to its target. This givesHST the ability to remain pointed at that target with no more than 0.007arcsec of deviation over long periods of time. This level of stabilityand precision is the equivalent of being able to hold a laser beamfocused on a dime 200 miles away for 24 hours.

Additionally, FGS provide precise astrometrical measurements of starsand celestial objects, which are advancing the knowledge of stars'distances, masses and motions. FGS provide star positions that are aboutten times more accurate than those observed from ground-basedtelescopes. When used as science instruments, the sensors allow HST tosearch for a “wobble” in the motion of nearby stars, which may indicatethat they have planets orbiting around them; determine if certain starsare actually double stars; measure the angular diameter of stars andother celestial objects; refine the positions and the absolute magnitude(brightness) scale for stars; and help determine the true distance scalefor the universe.

All telescopes have optical systems, and some even have specializedinstruments, but HST is almost unique in that it operates in space; thetelescope is actually “flown” as a spacecraft. Therefore, several spacecraft systems are required to keep HST functioning smoothly. Theessential systems are communications antennae, solar arrays for power,computers and automation, and housing.

HST performs only in response to detailed instructions from aground-based control center, and thus communications antennae arenecessary to transmit and receive such instructions between thetelescope and the Flight Operations Team at the Space Telescope ScienceInstitute. The four antennae on HST transmit and receive data via one ofthe constellation of Tracking and Data Relay Satellites (TDRS) operatedby NASA. In order for this system to be operational, at least one TDRSsatellite must be visible within HST's line of sight. Direct interactioncan occur between HST and the control center only when this line ofsight exists. When none of the TDRS satellites are visible from HST, arecorder stores the accumulated data until visibility is resumed. A flowdiagram of the communications process is provided as FIG. 4.

Flanking HST's tube are two thin, blue solar panel arrays. Eachwing-like array has a solar cell “blanket” that converts the sun'senergy directly into electricity to power HST's various systems. Some ofthe energy generated by the arrays is stored in onboard batteries sothat HST can operate while traveling through earth's shadow (about 36minutes out of each 97 minute orbit). Fully charged, each batterycontains enough energy to sustain HST in normal science operations modefor 7.5 hours, or five orbits. The solar arrays are designed forreplacement by visiting astronauts aboard a space shuttle.

In order to run all the many subsystems onboard HST, several computersand microprocessors reside in the body of HST, as well as in eachscience instrument. Two main computers, which girdle HST's “waist,”direct all operations. One communicates with the instruments, receivestheir data and telemetry, sends the data to interface units fortransmission to the ground, and sends commands and timing information tothe instruments. The other main computer handles the gyroscopes, thepointing control sub-system, and other HST-wide functions. Eachinstrument itself also houses small computers and microprocessors thatdirect their activities. These computers direct the rotation of thefilter wheels, open and close exposure shutters, maintain thetemperature of the instruments, collect data, and communicate with themain computers.

In space, HST is subject to the harsh environment of zero gravity andtemperature extremes—more than 100 degrees Fahrenheit difference intemperature during each trip around earth. To accommodate this operatingenvironment, HST has a “skin,” or blanket, of multilayered insulation,which protects the telescope during temperature shifts. Beneath thisinsulation is a lightweight aluminum shell, which provides an externalstructure for the spacecraft and houses the OTA and science instruments.The OTA is held together by a cylindrical truss made of graphite epoxy,the same material used to make many golf clubs, tennis racquets andbicycles. Graphite is a stiff, strong and lightweight material thatresists expanding and contracting in extreme temperatures.

The following table summarizes some of the relevant facts about HST:

TABLE 1 Weight 24,500 Lbs. Length 43.5 Ft. Diameter 14 Ft. (Aft Shroud)Optical System Ritchey-Chretien Design Cassegrain Telescope PrimaryMirror 94.5 Inch Dia. Pointing Accuracy 0.007 Arcsec for 24 HoursMagnitude Range 5 Meters to 30 Meters (Visual Magnitude) WavelengthRange 1,100 Angstroms to 24,000 Angstroms Angular Resolution 0.1 Arcsecat 6328 Angstroms Orbit 320 Nautical Miles Inclined at 28.5 DegreesOrbit Time 97 Minutes per Orbit

As indicated by the foregoing narrative, HST was designed to be servicedand upgraded periodically throughout its lifetime. Specifically, asillustrated in FIGS. 5 and 6, the space shuttle program had plannedmissions dedicated to servicing and upgrading HST scheduled until 2010,at which time HST would be retired and de-orbited in favor of the newJames Webb Space Telescope (JWST). However, in January of 2004, NASAAdministrator Sean O'Keefe announced that he was canceling shuttleService Mission 4 (SM4) because of safety issues identified by theColumbia Accident Investigation Board (CAIB) Report. The CAIB Report wasissued as a result of the Space Shuttle Columbia disaster of Feb. 1,2003, when all seven astronauts aboard Columbia were tragically killedon reentry into the atmosphere.

Administrator O'Keefe's announcement presented the scientific communitywith at least two problems. First, without SM4, regularly scheduledupgrades to HST's scientific instruments could not be made, therebyconfining scientific use of HST to out-dated technology. Second, and ofgraver concern, without SM4, wearable parts on HST could not bereplaced. More specifically, gyroscopes necessary for HST's PointingControl System (PCS) are degrading and will probably cease to operatewithin the next three to five years, as indicated by the chart of FIG.7. The current PCS requires sensing from three gyros, and already onlyfour of the six gyros aboard HST are operational. Best estimates showthat a less than 50 percent probability exists that three gyros will beoperational by the late 2005 or early 2006. Although scientists aredeveloping a two-gyro pointing system, that solution may add only 12 to18 months of additional life.

Moreover, batteries that power HST's computers, instruments andvirtually all vital systems are at risk. Based on recent tests, each ofHST's six batteries is losing its charging capacity at a rate of 5.9 amphours per year. This is far higher than previous tested loss rates ofabout two amp hours per year, and points out a tendency for capacityloss rate to accelerate nonlinearly over time. Without intervention, asscheduled by SM4, it is projected that by the end of 2005 scienceoperations will likely require block scheduling and a lowering of thesafemode trigger. By 2006, it is probable that the state-of-charge atthe end of orbit night will be near the hardware sunpoint safemodetrigger, which is the lowest level of safemode that protects thevehicle.

Accordingly, there is a need for a method and attendant apparatus forautonomously servicing HST and other free-flying satellites duringflight using robotics.

BRIEF DESCRIPTION OF THE INVENTION

The problems and shortcomings experienced in the prior art, as detailedabove and otherwise, are substantially alleviated by this invention, aswill become apparent in the following recitation of the objects andbrief description thereof.

One object of this invention is to provide servicing of satellites andother spacecraft without human presence on or near the spacecraft beingserviced.

Another object is use the principle of supervised autonomy to controlone or more aspects of a spacecraft servicing mission.

Still another object is to provide a method of de-orbiting a spacecraftthat was not originally designed for de-orbit.

Accordingly, one method of the present invention comprises the nominalsteps of autonomously establishing a link between a spacecraft needingservicing and a servicing vehicle, and sending commands to a robotsystem aboard the servicing vehicle, such commands addressing theservicing needs of the spacecraft. The commands may be pre-scripted orgenerated on board the servicing vehicle by using embedded ArtificialIntelligence (AI) logic or other on board processing in response tosensed conditions. The robot system preferably comprises a grappling armand a dexterous robot, which, in combination, provide for most servicingneeds, such as replacing degraded batteries and upgrading the scientificcapabilities of the spacecraft by replacing scientific instruments. Theservicing vehicle may aid in other ways as well, such as by providingstowage area for replacement parts, by shielding the spacecraft from theharmful effects of the sun during servicing, and by boosting thealtitude of the spacecraft's orbit by firing its thrusters.

Of course, one skilled in the art will appreciate that executing thecommands effectuates the servicing procedure. Preferably, other suchcommands are also used and may include capturing the spacecraft to beserviced while still in free-flight by the robot system and providingthe spacecraft with an ability to de-orbit.

Under the principle of supervised autonomy, the commands sent to therobot system can be overridden by a ground-based operator, who is intelecommunication with the servicing vehicle. In one embodiment, theremote operator may be ground based. In satellite. Use of modular designand construction of satellites and commonality among satellite designscan be used to maximize the potential repairability of futuresatellites. Constellations of substantially identical satellites, suchas Iridium, GlobalStar and other communications satellites, stand togain the greatest potential benefit of the repair capabilities of thepresent invention due to the use of common parts and their relativeaccessibility to repair vehicles due to their relatively low earthorbit.

Even for satellites in higher orbits, such as geostationary orbits, itmay be more cost effective to conduct a robotic repair/servicing missionto extend the useful life of valuable communications and other satellitepayloads presently on orbit, rather than to construct and launch entirereplacement satellites. The use of the present invention also has thepotential to reduce the amount of orbital debris in orbit by removinginoperative satellites from orbit.

In the geostationary band, it is generally considered cost-prohibitiveto deorbit spent satellites, and satellites are generally boosted to ahigher orbit that will not affect satellites in the geostationary band.Use of a repair mission in accordance with the present inventionpotentially could reduce the problem of clutter from spent satellitesabove geostationary orbit by extending the useful life of satellites anddelaying their eventual boost to higher orbit. Even if it is necessaryto boost the ejection module into higher-than-geostationary orbit, onesuch module may be used to service multiple geostationary satellites,resulting in a lower net contribution to the clutter in thehigher-than-geostationary band.

One cause of end-of-life, particularly in geostationary satellites, isexhaustion of fuel used for stationkeeping and other maneuvers.Attaching a deorbit module to a geostationary satellite to provide forend-of-life boost above geostationary orbit can permit the use of fuelonboard the geostationary satellite for revenue producing purposes thatwas previously reserved (as required by international law andconventions) for the end-of-life boost maneuver. Exhaustion of fuel forattitude control thrusters may also cause a satellite to lose attitudecontrol capability and thus reach the end of its useful life. Inaccordance with the present invention, it is possible for the servicingvehicle to transfer additional fuel to the target satellite forstationkeeping in the geostationary band or continued use of attitudecontrol thrusters to allow it to extend its useful life.

The invention preferably also includes a method for capturing afree-flying spacecraft with a second free-flying spacecraft, the secondspacecraft including the autonomous operation abilities alreadydescribed. This method includes the steps of autonomously identifyingand pursuing the second spacecraft and then securing the two spacecrafttogether. Preferably, the second spacecraft has a grappling arm or otherdevice that aids in attaching the two spacecraft together, while thefirst spacecraft may have affixed berthing pins that facilitateattachment. One of skill in the art will recognize that a variety ofdocking mechanisms can be employed without departing from the scope ofthe present invention.

One skilled in the art may realize that the steps of autonomousidentification and pursuit might be carried out by comparing file imagesof the first spacecraft with real time images taken continuously by thesecond spacecraft. Of course, in this situation the second spacecraftincludes appropriate equipment to create images.

One method for capturing the spacecraft further includes the steps oflaunching the second spacecraft into space and establishing a linkbetween the second spacecraft and a remote operator. Alternately, asecond spacecraft previously launched and optionally used to conduct aprior servicing/de-orbit mission can be directed to initiate pursuit,capture, and repair/deorbit of a target satellite, either from a parkingorbit or directly from a first servicing rendezvous. In either case, theoperator can use a communication link to the second spacecraft tooverride the autonomous identification and manual pursuit, and manuallycontrol the entire procedure or any part thereof as desired. Similarly,the step of securing the two spacecraft together can be done eitherautonomously or manually. If AI logic onboard the second spacecraft isused to conduct the autonomous pursuit, capture, and repair/de-orbit ofthe target satellite, the ground operator preferably may take manualcontrol at any time either at his discretion or based on predeterminedcriteria. In addition, the second spacecraft may propose a particularaction and await approval from the ground operator to execute thataction depending on the inputs received by the AI logic or Authority toProceed conditions either preset in the AI logic or directed by theground operator.

Another method of this invention includes the steps of securing ade-orbit module to the spacecraft to be serviced and commanding thede-orbit module to de-orbit the spacecraft. For the purposes of thepresent invention, de-orbit includes both the removal of a spacecraftfrom orbit through reentry or transfer of a satellite to a disposalorbit, i.e. transferring a geostationary satellite to higher thangeostationary orbit for disposal at end of life. One skilled in the artwill appreciate that the de-orbit module allows for the flight path ofthe spacecraft to be controlled so that population centers can beavoided and impact can be targeted for an ocean or other desiredlocation. If it is desired to de-orbit multiple satellites, a pluralityof de-orbit modules can be transported by a single ejection module.Through sequential rendezvous and attachment of a de-orbit modulefollowed by detachment of the ejection module, a plurality of satellitesnot originally designed for de-orbit can be provided with a de-orbitcapability. Alternately, a satellite that has suffered a kick motorfailure that prevented it from reaching its desired orbit similarly canhave the faulty kick motor replaced with an operational replacement topermit the affected satellite to reach its design orbit.

DESCRIPTION OF THE DRAWINGS

In order that the claimed invention may be better made and used by thoseskilled in the art, and that the best mode of carrying out the inventionmay be more fully appreciated, the following drawings are provided, inwhich:

FIG. 1 is a cutaway three-dimensional perspective view of the HubbleSpace Telescope (HST), illustrating major components and systems.

FIG. 2 is a cutaway three-dimensional perspective view of the OpticalTelescope Assembly (OTA) of the HST.

FIG. 3 is a side view of the OTA illustrating major components and theoperation of the Ritchie-Chretien Cassegrain telescope design.

FIG. 4 is a schematic diagram showing the flow of data between HST andremote control centers.

FIG. 5 is time chart showing the scheduled HST servicing missions by thespace shuttle.

FIG. 6 is a three-dimensional perspective representation of the spaceshuttle servicing HST.

FIG. 7 is graph illustrating the projected deterioration of HST'sgyroscopes.

FIG. 8 is a bottom end view of the aft bulkhead of HST illustrating theberthing pins to which the Hubble Robotic Vehicle (RV) attaches atberthing.

FIG. 9 is a schematic of HST in alignment with the RV immediately priorto berthing, illustrating a preferred capture box and grapple fixturefor initial connection of the RV to HST.

FIG. 10 is a close-up view of the grapple fixture of FIG. 9.

FIG. 11 is a three-dimensional perspective view of the RV connected toHST, and shows the robotic servicing concept using a preferred grapplearm.

FIG. 12 is an isolated three-dimensional view of the De-orbiting Module(DM).

FIG. 13 is a three-dimensional perspective view of the Ejection Module(EM), illustrating the solar array panels in their open position.

FIG. 14 is a three-dimensional perspective view of the Grapple Arm (GA)shown in the extended position.

FIG. 15 is a three-dimensional perspective view of the Dexterous Robot(DR) shown with one arm in the extended position.

FIG. 16 is a schematic representation of the “Mission Manager” concept.

FIG. 17A is a chart summarizing the launch phase of a preferred methodaccording to the present invention.

FIG. 17B is a graphic representation showing a typical ascent for anAtlas V rocket, which could be used in conjunction with this invention,embarking on a Low Earth Orbit (LEO) mission.

FIG. 18 is a chart summarizing the commissioning phase of an embodimentaccording to the present invention.

FIG. 19 is a chart summarizing the pursuit phase of an embodimentaccording to the present invention.

FIG. 20 is a chart summarizing the proximity operations phase of anembodiment according to the present invention.

FIG. 21 is a chart summarizing the approach and capture phase of anembodiment according to the present invention.

FIG. 22A is a diagram representing the safety ellipse of the approachand capture maneuver profile of an embodiment according to the presentinvention.

FIG. 22B is a diagram representing the approach range of the approachand capture profile of an embodiment according to the present invention.

FIG. 23 is a chart summarizing the servicing phase of an embodimentaccording to the present invention.

FIG. 24 is a top view of the DR shown after coarse positioning iscomplete.

FIG. 25 is a schematic representation of a decision tree for determiningwhether a task is “robot-friendly.”

FIG. 26A is a side view of one end of a DR arm and a proposed target,illustrating the coordinate frames before calibration.

FIG. 26B is a side view of one end of a DR arm and a proposed target,illustrating the coordinate frames after calibration.

FIG. 26C is a side view of one end of a DR arm and a proposed target,illustrating the end of the arm in a hover position after correction.

FIG. 27 is a chart summarizing the ejection and disposal phase of anembodiment according to the present invention.

FIGS. 28A through 28D are three-dimensional perspective views of the DMillustrating sequential deployment of its solar arrays after separationof the EM.

FIG. 29 is a chart summarizing the science operations phase of anembodiment according to the present invention.

FIG. 30 is a chart summarizing the HST/DM disposal or de-orbiting phaseof an embodiment according to the present invention.

DETAILED DESCRIPTION

Reference is now made to the drawings, in which like numbers are used todesignate like features throughout. Some of the preferred objectives ofthis invention are to robotically service a target satellite, such asthe Hubble Space Telescope (HST) 50, and then provide HST 50 with ade-orbit capability, which will be used upon the completion of the HSTmission life. While this specification is describing the invention as itrelates to HST 50 as an example, one skilled in the art will recognizethat the methods and apparatus disclosed herein are given only asexemplary embodiments. Other embodiments exist that fit within the scopeof the invention as defined by the appended claims. Thus, for example,one skilled in the art will readily see that this invention hasapplication to any number of existing free-flying satellites needingservicing and/or de-orbiting, and not just to HST 50.

One embodiment of this invention comprises the Robotic Vehicle (RV) 100,which preferably is launched on an Expendable Launch Vehicle (ELV) 52.The RV 100 preferably includes three basic components: a De-orbit Module(DM) 102, an Ejection Module (EM) 104, and a Robot System (RS) 106contained within the EM 104. Whereas previous HST servicing missionsemployed astronauts, the method of this invention is conducted usingrobotics. After the completion of the servicing mission, the EM 104 andRS 106 preferably are released from HST 50 and optionally de-orbited fora targeted impact into the Pacific Ocean. Alternatively, the released EM104 and RS 106 may be placed in a parking orbit until needed forservicing another satellite. In contrast, the DM 102 preferably staysattached to HST 50 and eventually provides the de-orbit capability atthe end of the HST science mission.

As indicated, the RV 100 preferably includes the DM 102 and EM 104spacecraft, together with the RS 106. Preferably, a two-fault tolerantseparation clamp band (not shown) releasably joins the DM 102 and EM 104segments mechanically, while electrical harnesses support variousCommand & Data Handling (C&DH), Guidance, Navigation & Control (GN&C),and Electrical Power System (EPS) functions across the interface. The RV100 is preferably launched in this combined state and remains combineduntil the completion of the servicing phase of the mission, at whichtime the EM 104 separates from the DM 102, while the DM 102 remainsattached to HST 50 through end-of-mission. One of skill in the art willrecognize that a variety of mechanism for separating DM 102 and EM 104may be employed without deviating from the scope of the presentinvention. In addition, the combination of DM 102 and EM 104 into asingle spacecraft is also within the scope of the present invention.

In one embodiment of the combined vehicle configuration, the DM 102provides the “brains” and the EM 104 provides the “muscle,” with respectto overall vehicle control. The relative navigation sensors and thealgorithms to determine absolute RV 100 attitude and relative attitude(RV 100 to HST 50) reside in DM 102. The GN&C system actuators,thrusters and momentum management devices reside in the EM 104 andrespond to the DM 102 commands to obtain the required RV 100 vehiclestate. The respective locations of the sensors, actuators, and momentummanagement devices can be redistributed between the two vehicles in aplurality of combinations without deviating from the scope of thepresent invention.

In one embodiment, the primary functions of the DM 102 are to provideGN&C attitude determination and control functions for the RV 100,mechanical and electrical interfaces to HST 50, mechanical andelectrical interfaces to the EM 104, and controlled reentry of the DM102 HST 50 combined vehicle at end-of-mission

The DM 102 GN&C subsystem preferably includes relative and absoluteattitude sensors and their associated algorithms, and actuation hardwareto provide three-axis stabilized attitude control. The preferred sensorsuite comprises sensors that are both independent and overlap incoverage (range to target) capability, and provides data used by thecontrol algorithms to determine the HST 50 relative state (ororientation) during the automated rendezvous and capture tasks, asdescribed hereafter. Other sensors provide absolute attitude informationto maintain desired vehicle pointing during all mission phases. Thissystem, when combined with the actuators in the EM 104, brings the RV100 into a position that facilitates either a first embodiment whereinthe capture of HST 50 is accomplished using the RS 106 Grapple Arm (GA)108 to capture one of the HST 50 grapple fixtures (see FIGS. 9 and 10)or a second embodiment wherein direct docking to the HST 50 berthingpins located on the aft bulkhead (see FIG. 8) is used. The system allowsfor capture even in the case where HST 50 is unable to maintain anotherembodiment, the remote operator may be space based, i.e. resident on theSpace Shuttle, International Space Station, or other human occupiedspacecraft. In yet another embodiment, the operator may be based on aseaborne platform or airborne platform. Thus, any command that may needto be changed can be done with little or no interruption of theservicing procedure, or, alternatively, the operator can manually carryout the servicing procedure without using the autonomous commands. It isalso preferred that one or more of the pre-scripted commands requirevalidation by the ground-based operator prior to it being carried out,thereby providing a stop-check against accident or error.

The preferred apparatus to provide a de-orbiting capability to thespacecraft includes a de-orbit module, which has thrusters and aguidance, navigation and control system for directing the flight path ofthe spacecraft as it de-orbits. Moreover, the apparatus also may includean ejection module, which contains the robot system, as well as partsfor servicing the spacecraft. One skilled in the art will understandthat the de-orbit module and the ejection module may separate afterservicing is complete, with the de-orbit module staying with theserviced spacecraft until it is ready for de-orbit at a future time.Meanwhile, the ejection module, after separation from the de-orbitmodule, may de-orbit itself immediately, or, alternately, may proceed tocapture another spacecraft in need of servicing. The ejection moduleoptionally may be placed in a parking orbit for some period of time toawait instructions to rendezvous with and service and/or deorbit anotherspacecraft, either at a predetermined time or in response to someunforeseen contingency or component failure on the other spacecraft inorbit.

Tools and space parts for a plurality of spacecraft may be transportedon the ejection module to facilitate servicing and/or de-orbit of aplurality of satellites. Such a generic or multi-satellite servingvehicle may be of particular interest to manufacturers of families ofsatellites using multiple substantially similar satellites with commoncomponents. Such a vehicle could potentially reduce or eliminate theneed for costly on-orbit or ground based spare satellites. Rather thanplacing one or more on orbit spares in orbit in anticipation ofpotential failures, the present invention provides an alternative meansto provide a flexible servicing capability that may enable rapid returnto service of a variety of satellites without requirement for entirededicated spares for each type of a controlled attitude due todegradation of its pointing and control system, which occurs as a resultof failing gyroscopes. The GN&C subsystem also supports the controlledHST reentry task at end-of-mission.

The relative navigation sensor selection is based on requirements forredundancy, range capabilities, and Technology Readiness Level (TRL).Two different types of sensors are required for redundancy. Thefollowing exemplary sensors may be used under this criteria, althoughone skilled in the art will understand that other sensors andcombinations of sensors fall within the scope of this invention:

Longer Range Sensors (Beginning at 5-3 km):

-   -   Primary: Optech Light Detection and Ranging (LIDAR)—Manufactured        by MDR    -   Secondary: Laser Camera System (LCS)—Manufactured by NEPTEC        Close Range Sensors (10 m and closer):    -   Primary: Enhanced Advanced Video Guidance Sensor (EAVGS)    -   Secondary: Natural Feature Imaging Recognition (NFIR)—System of        Eight Digital Video Cameras with Various Focal Lengths        Specifically Positioned to Align with Various HST        Features/Targets

The following table summarizes the sensor effective ranges for theexemplary sensors relative to HST 50. The dark grey shading shows theeffective range of the selected sensor. The light grey shading for theLCS shows the extended range for secondary preferred sensors.

TABLE 2

The DM 102 preferably includes a propulsion subsystem, which comprises anumber of Small Reaction Control System (RCS) thrusters coupled to fourlarge primary thrusters, a plurality of propellant tanks, and associatedvalves, filters, and lines. Monopropellant (high purity) propellant ispreferably used to fuel the thrusters in order to minimize plumecontamination. The DM 102 propulsion subsystem preferably is only usedfor the DM 102/HST 50 controlled reentry mission phase (describedhereafter) in order to minimize propellant slosh during attached scienceoperations. Thus, the DM 102 propulsion subsystem preferably remainspressurized and sealed until actual use in de-orbiting HST 50. At thattime, the RCS thrusters are used to point the combined vehicle to thecorrect attitude in response to the GN&C commands and the primarythrusters will perform the large delta-v burns required for controlledreentry. One of skill in the art will recognize that the foregoingpropulsion system is merely exemplary and that a variety of propulsiondesigns and components alternatively, including, but not limited to,cold gas thrusters and bipropellant liquid rocket thrusters, may be usedwithout deviating from the scope of the present invention.

The C&DH provides the functions required for acquiring, processing,storing, and transferring commands and telemetry. More specifically, theDM 102 C&DH also performs integrated operations with the EM 104 and asafe hold computer. The next table provides a breakdown of an embodimentof a C&DH control and operation for one procedure or method of thisinvention. All functions, with the exception of input/output to somespecialized equipment, preferably utilize a MIL-STD-1553 for exchange ofdata and commands. The preferred processor is a RAD750 on a CompactPeripheral Component Interface (CPCI) Bus based on a heritage processordesign used on the Mars Reconnaissance Orbiter program. Also, a VIRTEXII based imaging processing platform is preferably used for simultaneousprocessing of multiple (at least five) video sensors. Hot redundantoperation of the system preferably provides for concurrent processing.

TABLE 3 EM Mission RV DM C&DH C&DH DM EM Phase Control Mode Mode RIU ACELaunch ELV/DM Idle to Idle Disabled Enabled Safe (Breakwire- based)Commissioning DM Normal Idle Disabled Enabled Pursuit DM Normal IdleDisabled Enabled Prox Ops DM Normal Idle Disabled Enabled Approach & DMNormal Idle Disabled Disabled Capture (Warm- (Extended Backup) Time-out)Servicing DM Normal Idle Disabled Enabled EM Jettison & EM Standby Free-Enabled Enabled Deorbit Flight Science HST/DM Standby NA Disabled NA(Safemode) DM/HST DM Normal NA Enabled NA Deorbit

The DM 102 communications subsystem is preferably made up of an S-bandsystem using two multi-mode transceivers routed to a pair of low-gainantennas (LGA) for command uplink and telemetry transmission at low datarates through the Tracking and Data Relay Satellite (TDRS) system.High-rate DM 102 data, e.g. video, required during pursuit, capture,proximity operations, and servicing phases is passed through the EM 104communications system (described hereafter) Ku-band transmitters andhigh-gain antennas for downlink through the TDRS.

The preferred DM 102 S-band system through the TDRS is capable ofcommand reception at 2 Kbps and 16 Kbps, and transmits telemetry atbetween 2 and 16 Kbps. Similarly, the preferred DM 102 S-band systemrouted through EM104 will be capable of command reception at 2 Kbps andtransmit between 4 and 16 Kbps.

As one skilled in the art will see, the DM 102 mechanical subsystemdesign is driven by requirements of the launch and science operationsmission phases. During the launch phase, the DM 102 structure preferablyis connected to and interfaces with the Expendable Launch Vehicle (ELV)52 Payload Adapter Fitting (PAF) 54 on one side and to the DM 102/EM 104separation ring 110 on the other. In this configuration, the DM 102 hasto support the EM 104 mass during the launch phase and meet primary moderequirements levied by the ELV 52 provider. Requirements for the scienceoperations phase push the design to be as light and stiff as possible(preferably moments of inertia below 166,000 kg m2 and an approximately20 Hz first mode).

One embodiment of a structural design that meets these requirements isaluminum honeycomb forward and aft panels 112 with a segmented centerbulkhead 114 attached to a central tube 116 for internal frameattachment. FIG. 12 presents an embodiment of a structural design with atypical component layout.

The DM 102 electrical power system (EPS) provides power for the DM 102during all flight phases and power augmentation to HST post-servicing.One embodiment includes about 90 ft² of effective solar array area usingtriple-junction GaAs cells, ten 55 amp hour Li-ion batteries, andassociated power conditioning and distribution hardware known to oneskilled in the art.

In one embodiment, the DM 102 thermal control subsystem must controltemperatures within hardware limits and insure that there is no greaterthan about 5 W of thermal conductivity from the DM 102 and HST 50 duringattached science operations. One embodiment, based on prior known andproven methodology, utilizes heat pipes and software controlled heatercircuits to provide the required hardware component thermal environmentsfor all mission phases. Thermal isolation of the berthing latches orpins 118 minimizes heat conducted to HST during science operations.

Turning attention now to the EM 104, best seen in FIG. 13, in oneembodiment, EM 104 preferably functions to provide and/or perform:

-   -   GN&C actuation for the RV 100.    -   Provide all propulsion during pursuit, rendezvous, capture and        servicing of HST 50.    -   Equipment stowage for items that will be replaced on HST 50 as        well as for the removed equipment.    -   Mechanical and electrical interfaces to the RS 106.    -   Mechanical and electrical interfaces to the DM 102.    -   Controlled reentry of itself and/or rendezvous with another        satellite.    -   Provide high-rate Ku-band high-gain-antenna communication        system.    -   Provide shadowing for instruments during servicing.    -   Provide solar array (SA) for electrical power when charging HST        50 during servicing.    -   Provide reaction wheel assembly (RWA) control.

The EM 104 GN&C subsystem preferably comprises primarily actuators, suchas momentum and/or reaction wheels, magnetic torquers, and thrusters, toenable RV 100 pointing, and sensors, such as coarse sun sensors, aplurality (preferably three or more) of axis magnetometers, IMUs andGPS, to enable three-axis stabilized attitude control during EM 104controlled reentry. The relative navigation sensors and the algorithmsto determine absolute RV 100 attitude and relative attitude (RV 100 toHST 50) reside in the DM 102, as explained above. During the majority ofthe mission, the EM 104 GN&C system actuators, thrusters and momentummanagement devices preferably respond to the DM 102 commands to obtainthe required RV 100 vehicle attitude, i.e. the DM 102 controls the EM104/DM 102 stack. However, the EM 104 Actuator Control Electronics (ACE)preferably controls its own propulsion system and the EM 104 providesthruster and momentum/reaction wheel actuation throughout all of itsprimary mission phases. RV 100 attitude control may be provided by theDM 102 during pursuit, proximity operations, capture, docking andservicing. The DM 102 preferably also controls the attitude of both RV100 and HST 50 throughout the servicing phase of the mission, exceptwhen the EM 104 enters into a contingency safehold status. This safeholdstatus preferably is entered if communication is lost between the DM 102and the EM 104, but may also be entered if another unsafe condition isdetected or if commanded by the ground operator. Further, the EM 104 ACEsoftware can provide nominal attitude control during EM 104 separationand de-orbit. In order to establish positive control, the EM 104computer will assume attitude control of the RV 100/HST 50 stack justprior to separation.

An embodiment of the EM 104 propulsion subsystem includes a number ofRCS thrusters coupled to four large primary thrusters. In one embodimentof the invention, the EM 104 propulsion system is only used duringmission phases involving the RV 100 as a whole, and, after separation,is used for EM 104 controlled reentry and/or rendezvous with anothersatellite. One skilled in the art will appreciate that the RCS thrustersand momentum management hardware can point the RV 100 to the correctattitude in response to the GN&C commands received from the DM 102 andthe primary thrusters will perform the large delta-v burns required forrendezvous operations with HST 50. The RCS and momentum managementhardware will be used during proximity operations to bring the RV 100into position for HST 50 capture.

Post-capture, this propulsion system preferably is used to maintain theattitude of the RV 100/HST 50 combined vehicle. Upon completion ofservicing, the EM 104 may separate from the DM 102 and perform acontrolled reentry and/or second rendezvous using these RCS thrustersfor pointing and the primary thrusters for any large delta-v burns. Thecurrent preference is that EM 104 will carry 4800 lbs of hydrazine fueldistributed among five tanks. Thruster impingement will be controlled toreduce contamination and attitude disturbances to reasonable levelsestablished in the art. In another embodiment, the EM 104 propulsionsystem is made up of four 100 lb, four 20 lb and 36 7 lb thrusters. Oneof skill in the art will recognize that a wide variety of thruster sizesand combinations can alternately be used without deviating from thescope of the present invention.

Preferably, the EM 104 C&DH functions to collect and downlink telemetry,collect and downlink video, receive and process spacecraft and robotcommands and perform EM 104 separation and de-orbit tasks. The C&DHcommunication board (not shown) provides a preferred S-band uplink witha nominal 16 Kbps command rate and a 2 Kbps command rate contingencymode. The EM 104 C&DH in one embodiment can service up to 16 hardwarediscrete commands. The S-band downlink rates are preferably 2, 4, 8 and16 Kbps, where 8 Kbps is the nominal rate. Ku-band downlink ratepreferably is approximately 50 Mbps. The Ku-band downlink may beReed-Solomon encoded, although one skilled in the art will recognizethat other types of encoding may also be used without changing the scopeof this invention. Furthermore, all data preferably is delivered to theKu-band receiver within 200 ms of generation. This C&DH embodimentpreferably also provides a continuous command stream at a rate of 10 Hzto the RS 106. In one embodiment, the C&DH main components are a RAD750processor board, RS interface card, S-band communication card, Ku-bandcard and a low voltage power supply card.

The EM 104 communications system preferably contains an S-band systemusing two multi-mode transceivers routed to a pair of Low Gain Antennas(LGA) for command uplink and telemetry transmission at low data ratesthrough the TDRS system. In addition, the preferred system also containsa pair of Ku-band transmitters and steerable High Gain Antennas (HGA) totransmit real-time video and high-rate engineering telemetry throughTDRS.

The EM 104 S-band system through the TDRS system is preferably capableof command reception at 2 Kbps and 16 Kbps, and will transmit telemetryat between 2 and 16 Kbps, while the EM 104 Ku-band system preferablytransmits channeled video and telemetry at 50 Mbps and 128 Kbps,respectively. The EM 104 S-band system through the Ground Network/DeepSpace Network will be capable of command reception at 2 Kbps andtransmit telemetry at between 4 and 16 Kbps.

The mechanical subsystem of the EM 104 preferably includes four mainsections, namely structure, mechanisms, solar array, and high gainantenna systems. Preferably, structures include propulsion, avionics,and robotic and instrument modules. One skilled in the art willunderstand that the propulsion module houses propulsion tanks, valves,plumbing, and fill and drain valves, while the avionics module housesbatteries and electronics boxes, except for placement of critical GN&Csensors and actuators. Also, the majority of the harnesses preferablyare housed in this section. In one embodiment, the robot and instrumentmodule houses replacement instruments for HST 50, such as the Wide FieldCamera 3 (WFC3), Cosmic Origins Spectrograph (COS), and Fine GuidanceSensors (FGS), as well as the Dexterous Robot 120, the GA 108, conduits,tools, and tool caddies. Each of these components will be betterdescribed hereafter.

The mechanisms subsystem preferably includes the compartment doormechanisms and orbit replacement instrument (ORI) stowage bays. In oneembodiment, the WFC3, COS and FGS stowage bays are designed to provideadequate mechanical isolation during all phases of the RV 100 mission.The EM 104 solar array subsystem preferably provides solar shadowingprotection to the ORIs and exposed HST 50 cavities during transport andservicing activities. This is done while minimizing the shadowingimpingement onto the HST 50 solar arrays. One embodiment of the solararray for the EM 104 is a 14-string daisy-style solar array providingshadowing across the servicing workspace and recharging batteries forthe EM 104 during servicing activities. Ten Li-ion batteries arepreferred that provide EM 104 power of approximately 3700 W (payload isabout 3000 W) during servicing.

An embodiment of the HGA subsystem has an azimuth motor with 360 degreerotational freedom coupled with a +/−90 degree elevation motor whichallows the EM 104 HGAs complete hemispheric visibility.

In another embodiment, the EM's electrical power subsystem switches theprimary DC power to the EM 104 actuators, mechanisms, sensors, antennas,DR 120, Grapple Arm 108, and electronics boxes. One skilled in the artwill understand that the EM 104 electrical power system measures andreports voltages, currents and status of components connected to the EMpower bus. In one embodiment, electrical power may be provided foroperation of about 200 temperature sensors and approximately 22 heaters.

Referring more specifically to FIGS. 14 and 15, the RS 106 is nowdiscussed. Preferably, RS 106 is housed in the EM 104 and comprisesgenerally the GA 108 and the DR 120, along with appropriate controlelectronics, such as a vision system (VS), tools and tool caddies forservicing HST 50.

Preferably, the GA 108 is a multi-axis manipulator with a Grapple ArmEnd Effector (GAEE) 122 on one end providing a power/data/videointerface. One embodiment calls for the GA 108 to be responsible forcapturing HST 50 via either of the two grapple fixtures (in conjunctionwith GN&C system), as shown in FIGS. 9 and 10, and for positioning theDR 120 for servicing as required. This preferred GA 108 provides sixdegrees of freedom and a 39 ft. total reach (one 20 ft and one 19 ftlink). There are preferably two cameras 126 and 128 on the elbow 124 andtwo cameras 130 and 132 on the GAEE 122. The GA 108 of this embodimentweighs about 1500 lb and will operationally require approximately 210 W.peak, 100 W. average, power. One of skill in the art will recognize thatthe foregoing robot system is merely exemplary and a variety of roboticmechanisms and techniques known in the art could be substituted withoutdeparting the scope of the present invention.

One embodiment of DR 120, illustrated in FIG. 15, is responsible forservicing HST 50, in conjunction with the GA 108 and attendant tools. Itis comprised of dual manipulator arms 134 and 136 (about 11 ft totallength each) with 7 degrees of freedom each, multiple camera units withlights, and an Orbital Replacement Unit (ORU) Tool Change-out Mechanism(OTCM) 138 or end-effector (not shown) on the end. The DR 120 of oneembodiment weighs approximately 2950 lb. and will operationally require2000 W. peak, 1700 W. average, power. The size, weight, and powerrequirements for the DR and its robotic mechanisms may be varied toaccommodate servicing missions on satellites of various sizes andcomplexities

Preferably, the VS is responsible for and displays situational awarenessduring RV 100 mission phases. The VS provides an integrated camera andvideo delivery system comprising a series of cameras, lights, andelectronic boxes that reside at various locations on the RS 106 and theRV 100. The VS provides visual verification and active vision, usingstill and streaming video from its complement of cameras, feed back forautomated sequencing, manual and scripted servicing, visual inputs to AIlogic, and error model correction, and HST 100 inspection and surveyactivities.

One embodiment of a VS complement of cameras includes eight on the robottorso, eight on the OTCM 138, two LIDAR cameras, two on the GA 108 elbow124, two on the GAEE 122, two on the bay and about fourteen situationalawareness cameras for a total of 38 cameras. Those skilled in the artwill realize that each of these cameras is connected to a Video ControlUnit (VCU). The VCU collects the video data from the active camerasconnected to it and converts the raw image data, preferably to JPEG2000.The JPEG2000 format includes any compression, if selected.

Moreover, one embodiment of the VS contains a command and telemetryinterface to the EM 104 C&DH that enables VS control, configuration,state of health and video distribution. Remote commands may be used tocontrol power, frame rates, and compression ratios of each camera. Twoadditional stereoscopic cameras may be used to provide detailed worksiteviews and support machine vision function for worksite localization thatallow the remote operators to monitor robot configuration, servicingtask operations, and other activities

Looking now at the method of carrying out the invention, one skilled inthe art will readily observe that the preferred objectives of themethods for the exemplary HST servicing mission are threefold:

-   -   To enable the safe disposal of the HST 50 when it reaches the        end of its useful life.    -   To implement life extension measures that will assure HST 50        mission life for at least five years beyond the completion of        the servicing mission.    -   To enhance the scientific capability by installing the        scientific instruments planned for SM-4.

One embodiment of the invention meets these objectives as follows.Towards meeting the first objective, the DM 102 may to be attached tothe HST 50 for later use as a de-orbit vehicle. To meet the second, anew set of gyroscopes, batteries, and a replacement FGS may beinstalled. And, to meet the third objective, the WFC3 and COSinstruments may be installed. Each of these procedures, as well as apreferred over-all mission procedure, will be described in greaterdetail.

One embodiment of a Robotic Servicing and De-Orbiting Mission (RSDM)mission may be defined by the following phases or steps: launch, RVcommissioning, pursuit, proximity operations, approach and capture,servicing, EM ejection and disposal, science operations, and HST/DMdisposal. The description that follows will specifically discuss aservicing mission to HST 50. However, one of skill in the art willrecognize that the principles disclosed herein can be applied toservicing and/or de-orbiting of a variety of spacecraft in orbit aroundthe earth or elsewhere in space. It will also be clear to one of skillin the art that the robotic servicing and de-orbit phases may beimplemented independently; servicing without de-orbit or de-orbitwithout servicing may be implemented for a particular spacecraft ifdesired. If the target satellite is in an orbit, such as geostationary,where de-orbit may not be practical or cost effective, a de-orbit modulecan be used to place the target satellite into a disposal orbit, forexample above the geostationary band. The following table summarizeseach phase of the exemplary mission.

TABLE 4 Phase Duration Phase Begins Phase Ends Highlights Launch 2-3hours Launch RV in a power Launch, stage 1 positive ignition/separation,configuration stage 2 ignition/separation, RV separation, EM establishespower positive attitude with solar panels deployed. RV 14 days RV hasCheckout Checkout RV Commissioning established activities systems andGA. power positive complete. configuration. Pursuit 2-12 days RV Endsjust prior Includes orbital commissioning to maneuver to maneuvers tobring completed. the safety ellipse. the RV to the final co-ellipticorbit within relative navigation sensor range of the HST. Proximity 1-2days Begins with the Ends just prior Enter safety Operations maneuverfrom to leaving the ellipse, survey the final co- final safety HST, HSTand RV elliptic orbit ellipse. preparations for to the Safety capture.Ellipse. Approach and 2 hours RV maneuvers Ends with the Deploy GA,Capture away from completion of grapple HST, the safety the mechanicalposition HST in ellipse to berthing/docking berthing latches the captureof the RV to HST and partially close axis. with the mated latches,maneuver spacecraft (RV/HST) stack to sunpoint in a preferred sun-attitude, complete pointing attitude. latch closure. Servicing 30+ daysMechanical Ends when the EM Battery docking receives authorityaugmentation, complete. to take control of WFC3/RGA II, the HST/RV stackCOS, FGS, reboost prior to separation. (optional). EM Ejection & 4 daysEM takes Completion of Separate EM from Disposal control of controlledDM, perform HST/RV stack disposal evasive maneuvers, from the DM. of theEM. perform de-orbit burns, impact in Pacific. EM continues this phasewhile HST/DM enters science operations. Science 5+ years EM ejection Endof science Verification, Operations from HST/DM observations and nominalscience stack. end of life testing. operations. HST/DM 4 days End ofscience Safe disposal of De-orbit burns, disposal operations. HST/DM.impact.

In one embodiment, RV 100 and HST 50 each have their own safing systems.Accordingly, prior to initiation of the capture phase, the HST safingsystem is disabled to prevent any unintended change in attitude or solararray motion. The system will remain disabled until after the EM 104 hasseparated from the DM 102, although some HST 50 tests may be enabledduring the mission to protect against inadvertent solar array motion.The following table shows which vehicle has primary safing authority foreach mission phase in one embodiment.

TABLE 5 Vehicle Software Safing Hardware Configuration Mission PhaseStrategy Safehold RV Launch, Pursuit and DM Software Safe Hold, EMProximity Operations Mission Manager Abort RV Capture DM (UseRedundancy), EM Mission Manager Abort RV + HST Servicing DM SoftwareSafe Hold EM EM Ejection EM EM EM EM De-orbit EM EM DM + HST Science HSTwith DM backup HST PSEA DM SHM Disabled DM + HST HST De-orbit DMSoftware Safe Hold, DM Mission Manager Abort

Preferably, the DM 102 will utilize previously known Mission Manager(MM) software for most Guidance, Navigation and Control (GN&C)commanding requirements, although one skilled in the art will observethat other software can used to accomplish this function withoutdeparting from the scope of this invention. MM allows the RV 100 toimplement the GN&C sequences autonomously with optional remoteintervention programmed in at key points. FIG. 16 graphically depictsthe MM concept. As indicated, this software utilizes task lists that aretied to the mission phases as discussed in the following sections. MMmay hold up to five nominal and five abort task lists that are uploadedone at a time from the remote operator, each containing up to 100 tasksper list, although only one task list will be active at any given time.When received by MM, a remote-specified task list preferably isvalidated, and then moved to the specified task list slot. Each task ina nominal task list preferably has a corresponding abort task list slot.If an abort occurs during a task, the associated abort task list will beinitiated for the abort sequence. A task list preferably is not allowedto span mission phases, and a separate GN&C remote command is requiredin order to change a mission phase. Each task within a task list maycontain an Authority To Proceed (ATP) flag that, if set, will allow thattask to proceed to the subsequent task in the list upon nominalcompletion. If the ATP is not set (i.e. ATP flag=0), the task willsuspend upon nominal completion and await explicit authorization toproceed from the remote operator. If desired, a second task list may beset to “pending,” such that it will automatically execute pendingnominal completion of the active task list. Preferably, only one tasklist can be defined to be active, and only one task list can be definedto be pending at any given time.

Referring now to FIGS. 17A and 17B, the preferred launch phase starts atT-0 and terminates when the RV 100 achieves a power positiveconfiguration after separation from the second stage. FIG. 17A is asummary of the launch phase and FIG. 17B shows a typical ascent for anAtlas V expendable launch vehicle 52 engaged in a low earth orbit (LEO)mission. Based on the size of the RV 100, an Evolved Expendable LaunchVehicle (EELV) 52, i.e. an Atlas V or a Delta IV rocket with a fivemeter fairing could be used.

The major systems of the RV 100 preferably are configured for launch asfollows:

-   -   DM 102 C&DH prime on    -   EM 104 C&DH prime and redundant on    -   All receivers on    -   ACE box off    -   Reaction wheels off

The launch dispersions for the ELV 52 currently provide the followingpreferred accuracies: altitude ±10 km and inclination ±0.04°. In orderto further protect HST 50 from the RV 100 and to provide time forvehicle checkout after separation from the ELV 52 upper stage, thefollowing conditions are preferred in one embodiment:

-   -   Circular orbit 20 km below the altitude of HST 50 at the time of        launch.    -   Initial in-plane separation (phase angle) of 0-360 degrees,        which varies with launch epoch.    -   Right ascension difference to target HST 50 plane at end of        pursuit phase:        -   Allows relative drift of ascending node due to differential            gravitation (J₂).        -   A function of initial phase angle, semi-major axis            difference and time to rendezvous.        -   This drift constrains maximum time before first pursuit            phase maneuver.

The difference in semi-major axes will drive a differential rate ofregression of the Right Ascension of the Ascending Node (RAAN).Consequently, the target orbit may include a RAAN offset such that theorbit planes align at the nominal end of the pursuit phase. Thisdifferential RAAN constrains the allowable time before the firstmaneuver is executed; the first burn is nominally at launch +14 days.

Preferably, all activities in the launch phase are nominally to beperformed autonomously, meaning that each task is carried out withoutdirect human intervention. In one embodiment, each task may bepre-scripted. In another embodiment, at least some of the tasks may beinitiated by Artificial Intelligence (AI) logic resident in RV 100. Asin the case of pre-scripted commands, a capability for the remoteoperator to override any AI-generated commands may be provided.Alternately, the AI logic may propose a course of action that must beconcurred in by the remote operator before being executed. In yetanother embodiment, the remote operator pre-loads the GN&C launch tasklist into the DM 102, and after separation from the ELV 52, abreakwire—based DM 102 transition from idle mode to safe mode occurs.This also initiates the execution of the launch task list. The followingevents then may occur within the EM 102:

-   -   Turn on S-band transmitter    -   Turn on ACE    -   Power on the reaction wheels    -   Turn on the catalyst-bed heaters    -   Delay ten minutes    -   Open propulsion system valves    -   Activate thruster power bus    -   Start solar array deployment sequence (except the sections that        cover the stowed GA 108)

Next, the RV 100 preferably uses the reaction wheels to damp any tip offrates and then maneuvers to a −H1 sunpoint attitude. Torquer barspreferably are used to dump momentum. If the wheels saturate, the DM 102preferably activates momentum damping using thrusters, after allowingsufficient time for heaters to warm up the catalyst-beds. In anotherembodiment, a momentum wheel or wheels may be used to manage the tip offrates and to control the pointing attitude of RV 100.

RV commissioning preferably begins with the RV 100 in a power positiveattitude with the EM 104 daisy solar array deployed, and ends whencheckout of the RV 100 is complete and the maneuvers to initiate pursuitof the HST 50 are about to occur. FIG. 18 is a summary of the RVcommissioning phase.

Once on orbit, RV systems preferably are activated and verified forproper performance. Commanding comes from a combination of real-time andstored commands. In one embodiment, this process takes approximately 14days and is summarized as follows:

-   -   RV 100 orbit determination.    -   RV 100 GN&C checkout.    -   Activate Ku-band downlink.    -   Check out DM 102 sensors.    -   Activate LIDAR.    -   Activate and checkout GA 108.

The objective of these tests is to verify proper operation of the GA 108and measure performance that could not be verified directly on theremote station before it is used near HST 50. Each test includesaliveness, functional and zero-gravity performance tests. Preferably,the pursuit phase begins after the RV commissioning is completed andterminates just prior to the maneuver to place the RV 100 on the safetyellipse. All the GN&C activities preferably are contained in the pursuitphase task list and are performed autonomously with appropriateAuthority to Proceed (ATP) points for any actions that require a remotecommand to authorize continuation of the task list. FIG. 19 is a summaryof the pursuit phase.

The pursuit phase preferably includes a number of burns, detailed in thetable below. Preferably, the EM 104 provides the propulsion during thisphase using its RCS with the DM 102 providing the vehicle navigationcontrol. In one embodiment, the first two bums are executed to raise theco-elliptic orbit to 5 km below HST 50 and are executed after the 14 daycommissioning phase. This slows the relative in-plane drift between theRV 100 and HST 50. Initial sensor acquisition preferably happens at thispoint. As RV 100 closes on HST 50 from below and behind, two moremaneuvers preferably are made to raise the orbit to about 1 km below HST50. Again the relative in-plane drift is slowed. A maneuver may then bemade to correct for any out of plane error that may be due to launchdispersion or timing errors of the first two orbit boosts of the RV 100.The acquisition of HST 50 with relative navigation sensors such as LIDARmay now happen as RV 100 approaches within 5 km of HST 50, still frombelow and behind. After HST 50 acquisition, the RV 100 can now getnearly continuous range measurements to HST 50.

TABLE 6 Fuel Mass Delta V Consumed Maneuver MET Action m/s ft/s kg lbmContingency? 1 Launch plus Remove inclination error 5.3 17.4 37.9 83.6Launch dispersion 14 days (max 0.04 deg) Remove RAAN error from 2 2.27.2 15.7 34.6 Launch dispersion weeks inclination error Remove RAANerror from HST 0.8 2.6 5.7 12.6 Launch dispersion state error 2-3 Launchplus Remove altitude error from 5.5 18.0 39.2 86.4 Launch dispersion 14days launch 4-5 Launch plus Achieve coelliptic 5 km below 8.2 26.9 58.5128.9 14-30 days HST 6 Launch plus Remove remaining out of plane 0.5 1.63.5 7.8 16-31 days error 7-8 Launch plus Achieve coelliptic 1 km below2.2 7.2 15.6 34.3 17-33 HST  9-10 As required Achieve coelliptic 1 kmabove 1.1 3.6 7.8 17.2 Launch dispersion or HST 12.3 12.3 123.4 123.4failure to acquire rel-nav Total (no margin included) 25.8 84.6 184 405

The proximity operations phase preferably begins with the maneuver fromthe final co-elliptic orbit to the safety ellipse. And the proximityoperation phase ends just prior to the maneuver to leave the finalsafety ellipse to initiate the capture phase. Commanding of the RV 100in this phase is autonomous with ATP points. FIG. 20 is a summary of theproximity operations phase.

In one embodiment, as RV 100 passes below HST 50, the LIDAR acquires HST50 and a sequence of small maneuvers is performed in order to put RV 100on a 100 m×50 m×50 m Fehse-Naasz Walking Safety Ellipse (WSE) about HST50. The WSE is a natural relative motion, best described as a path alongthe surface of a cylinder with an axis along the HST 50 velocity vector.Consequently, the WSE is preferably strictly non-interfering with HST 50as the points where the RV 100 and HST 50 orbital planes intersectcoincide with maximum RV 100/HST 50 radial separation. The relativealong-track RV 100/HST 50 separation can be controlled by periodic smallmaneuvers (low delta-v). Another advantage of the WSE is that it allowsa thorough inspection of HST 50 as RV 100 circumnavigates it, therebyincreasing the ability to observe HST 50. If the LIDAR does not acquirethe HST 50, then the RV 100 passes safely under HST 50 and the maneuverplan will be revised. Burns for this phase are detailed in the followingtable:

TABLE 7 Fuel Mass Delta V Consumed Maneuver Duration Action m/s ft/s kglbm Contingency? 1-3 45 min Target safety ellipse 1 km 0.59 1.94 4.129.09 ahead of HST Midcourse correction 0.10 0.33 0.70 1.54 Achievesafety ellipse 1 km 0.39 1.28 2.72 6.01 ahead of HST as required Recoverto 1 km above or below 2.77 9.09 19.33 42.62 Failure to achieve safetyHST, approach HSt and repeat ellipse 1-3  4-11 18 hrs Maintain SE 1 kmahead for 0.03 0.11 0.23 0.51 18 hrs (TBR) 12-17 18 hrs Walking safetyellipse to 0.05 0.17 0.36 0.80 center HST as required 10 days Removerelative drift and 0.44 1.44 3.07 6.77 Provides 10 days on SE tomaintain HST-centered safety determine HST relative ellipse (every 3orbits for 10 attitude state days) Total (no margin included) 3.30 10.8130.54 50.70

In one embodiment, the approach and capture phase begins when RV 100leaves the WSE and terminates with the completion of the mechanicalberthing/docking of the RV 100 to HST 50 with the mated spacecraft in apreferred sun-pointing attitude. FIG. 21 is a summary of the approachand capture phase. Commanding of the RV 100 during this phase will beprimarily autonomous with ATP points programmed in.

In another embodiment, the approach and capture phase begins with asequence of maneuvers that puts the RV 100 on the capture axis of HST50. The capture axis is currently preferred to be the −V1, or in otherwords the negative end of HST's main longitudinal axis. In anotherembodiment, the −V3, or in other words negative lateral axis might beused since it has certain advantages in some situations (notably GS 108reach). The RV 100 preferably maintains its position on the capture axisat a fixed range (nominally 30 meters, but certainly outside thegeometry of HST 50) by matching rates with HST 50. Once the RV 100/HST50 relative rates are stabilized, and authority to proceed is granted,the RV 100 will descend down the capture axis to a stand-off distance˜1.7 m (for −V1 approach) from the capture target.

The preferred plan for the capture phase is dependent on the state ofHST 50 at the time of capture. If HST 50 is not functioning, ornon-cooperative, the nominal capture strategy will be as describedabove, with RV 100 approaching HST 50 along a potentially tumblingcapture axis. If HST 50 is functioning and cooperative, the nominalcapture strategy will be a traditional approach approximately along theR Bar (from the center of the earth outwards) to HST 50 in an inertialattitude configuration, with HST 50 pointing along the radial directionat the time of capture. This approach is accomplished by performing asequence of predominantly velocity direction maneuvers to allow RV 100to approach HST 50 from the nadir direction while minimizing thrusterplume contamination of HST 50. FIGS. 22A and 22B show a typical capturemaneuver profile. The two tables immediately below show the maneuverplans for a controlled and uncontrolled HST50, respectively.

TABLE 8 Fuel Mass Delta V Consumed Maneuver Duration Action m/s ft/s kglbm R-Bar approach 30 min Transfer along R-Bar to 30 0.56 1.84 3.88 8.55m-V1 hold from SE below HST LVLH Stationkeep 20 min 20 minute hold at 30m-V1 0.14 0.46 0.98 2.16 separation HST Stationkeep 10 min Approach to10 m hold point 0.10 0.33 0.67 1.48 HST Stationkeep 10 min 10 minutehold at 10 m hold point 0.05 0.16 0.52 1.15 HST Stationkeep 10 minApproach to 1 m hold point 0.04 0.13 0.37 0.82 HST Stationkeep 10 min 10minute hold at 1 m hold point 0.02 0.07 0.19 0.42 HST Stationkeep 10 minretreat to 10 meter hold point 0.08 0.26 0.78 1.72 HST Stationkeep 10min hold at 10 meter hold point for 0.16 0.52 1.56 3.44 10 minutes 2burn 45 min Safe return to safety ellipse 1.12 3.67 7.81 17.22 Total percapture attempt (no margin included) 2.27 7.45 16.76 36.95 Total for 4capture attempts (no margin included) 9.08 29.79 67.04 147.80

TABLE 9 Fuel Mass Delta V Consumed Maneuver Duration Action m/s ft/s kglbm Contingency? 2 burn 45 min Maneuver to predicted docking 1.50 4.9210.47 23.08 axis from SE HST 20 min 20 minute hold at 30 m-V1 4.00 13.1239.02 86.04 Stationkeep separation HST 10 min Approach to 10 m holdpoint 1.90 6.23 18.55 40.90 Stationkeep HST 10 min 10 minute hold at 10m hold point 0.90 2.95 8.79 19.38 Stationkeep HST 10 min Approach to 1 mhold point 0.40 1.31 3.90 8.60 Stationkeep HST 10 min 10 minute hold at1 m hold point 0.20 0.66 1.95 4.30 Stationkeep HST 10 min retreat to 10meter hold point 0.40 1.31 3.90 8.60 Failed capture Stationkeep HST 10min hold at 10 meter hold point for 10 0.90 2.95 8.78 19.36 Failedcapture Stationkeep minutes 2 burn 45 min Safe return to safety ellipse1.50 4.92 10.44 23.03 Failed capture Total per capture attempt (nomargin included) 11.70 38.39 105.82 233.29 Total for 4 capture attempts(no margin included) 46.80 153.54 423.28 933.17

At the hold point, the RV 100 preferably deploys the GA 108, as shown inFIG. 9, and captures either of the two grapple fixtures 140 located onthe −V3 side of HST 50, as shown in FIG. 10. The preferred candidatevehicle control configurations would have HST 50 in an inertial hold (ifthe PCS system is active) or tumbling, and the RV 100 going to freedrift once the GAEE 122 is inside a predefined capture box 142. Thecapture box 142 defines a set of conditions (position, orientation)within which the GA 108 can capture a grapple target. The GA 108preferably will be positioned so that a GAEE 122 camera 130 or 132 iscentered on the capture box 142. Capture is preferably planned so as tooccur during orbit night so the GA 108 lights can be used to control thelighting. Once the command to initiate the final capture sequence hasbeen given, the sequence preferably requires no operator intervention.However, manual override by ground control can always be done ifnecessary at any point.

In one embodiment, the RV 100 will notify the GA 108 when the grapplefixture 140 enters the specified capture box 142 with the requiredrates. The on-board vision software can then acquire the target andcommunicate its status to the RV 100. At that point the attitude controlmay be turned off to avoid inducing disturbance on the GA 108, and theGA 108 then moves to put the GAEE 122 on the grapple fixture 140, snarethe grapple or berthing pin 108, and rigidize. The GA 108 will absorbresidual relative vehicle rates. Subsequently, the GA 108 will apply itsbrakes and the RV 100 will null the combined vehicles rates to achieve apower positive, thermally stable, attitude profile. It may be necessaryto maneuver the stack to a −V1 sunpointing attitude to charge the RV 50batteries prior to berthing.

Once the rates have been nulled, the GA 108 will maneuver the HST 50 toa pre-defined pre-berth position relative to the RV 100. The operatorcan validate the script to maneuver to the berthing position at therobotics console using script rehearsal software, or can manuallymaneuver to the berthing position. Preferably, the operator then willbring the HST 50 berthing pins 118 into the capture mechanism on the DM102 using a combination of scripts and hand controller operations. Theoperator will then limp the GA 108 joints and allow the capturemechanism to complete the berthing sequence. Initially, the berthinglatches preferably are not completely closed. But after a predeterminedtime to allow the mechanism to equilibrate thermally, the latches may beclosed completely. In another embodiment, AI logic on RV 100 can executethe capture/berthing process with or without oversight and/orintervention from a remote operator.

An alternate preferred method of capture is to dock directly to the HST50 aft bulkhead at the berthing pins using a mechanism(s) on the RV 100.The candidate vehicle control configurations could have HST 50 in aninertial hold (if the PCS system is active) or tumbling at a rate of upto 0.22 deg/sec per axis and the RV 100 remaining in an active controlstate through capture.

The HST 50 could be in one of several attitude control modes for theRSDM, with the status of the Rate Gyro Assemblies (RGA) being theprimary determining factor. If three good RGAs remain, the HST 50 may beconducting normal science operations and the normal attitude control lawcan be used. In this scenario, science operations will be terminated atthe beginning of proximity operations so the HST 50 can be prepared forcapture and berthing. The science plan may include breakpoints at oneday intervals so that science observations can be extended if proximityoperations are delayed. In this case, the HGA booms will be retractedand the aperture door closed.

Given the history of the RGAs, it is quite possible that fewer thanthree healthy gyros will be available by the RSDM time frame. The HSTProject is developing alternate attitude control modes for science thatwill allow control on two or one gyro(s). Therefore, if one of thesemodes is in use at the time, the HST could terminate scienceobservations and remain in that control mode, either Magnetometer 2 Gyro(M2G) or Magnetometer 1 Gyro (M1G) mode. If no gyros remain, the HST 50preferably will be in safemode. However, if inadequate hardware remainson HST 50 to perform an attitude control function, the RV 100 can stillperform a capture either with the GA108 or by direct docking. One ofskill in the art will recognize that the above procedure can be modifiedto accommodate capture and service of three-axis stabilized, spinstabilized, or even satellites with non-functioning or partiallydisabled attitude control systems.

Preferably, the servicing phase begins once the RV 100 is completelyberthed to the HST 50, or other satellite to be serviced, and terminateswhen the EM 104 is given the control authority prior to EM 104 ejection.The hardware systems augmentation and science instrument change-outs mayoccur during this phase. During the servicing phase there preferably areno HST 50 maneuvers required, and thus, in a preferred embodiment, theRV 100 controls the combined vehicle attitude to maintain a powerpositive, thermally stable profile while accommodating any constraintsthat are imposed by the replacement instruments or apparatus, calledOrbit Replacement Units (ORU). FIG. 23 summarizes the servicing phase.

The preferred servicing tasks are planned to extend the life of HST 50,or other satellite to be serviced, and may provide enhanced science orother operational capabilities. To achieve these objectives the RS 106carries out a series of tasks, as described hereafter. Preferably, anynew hardware to be installed on HST 50 by the RS 106 is equipped withrobot-friendly interfaces that allow the ORU Tool Change-out Mechanism(OTCM) 138 end effecter to directly grasp and handle the ORU, includingthe new science instruments to be installed and interfaces on the RV 100to be actuated by the DR 120. However, the HST 50 or other satellite tobe serviced may not be equipped with the necessary handling interfacesfor the DR 120 to manipulate it directly. To address this problem, apreferred suite of specialized tools (not shown) has been developed tocreate the environment necessary for the RS 120 to carry out theservicing tasks on HST. One skilled in the art will recognize that otherspecialized tools could be used in conjunction with the servicing ofother satellites that would still fall within the purview of thisinvention. In one embodiment, the tools for use with the HST 50 includedevices to open the aft shroud, the radial bay, and bay doors; tools toactuate the latches to release the science instruments and remove theseinstruments from HST 50; and tools to mate and de-mate connectors,actuate bolts, and the like.

In addition to appropriate tools for the DR 120, successful completionof the preferred servicing objectives may require that the remoteoperators receive visual confirmation of the tasks being performed.Therefore, a vision system comprising multiple cameras with variousfunctions and specifications may be used to provide views of on boardactivities. In one embodiment, the preferred vision system, alreadydescribed above, is integrated with the DR 120, GA 108, EM 104 andservicing tools.

One skilled in the art will understand that due to the potential fordamage to detectors and thermal degradation of adhesives, the scienceinstruments, and the open cavities in which they are mounted, preferablyare protected from direct sun exposure. In one embodiment, thecombination of shading from the EM 104 solar arrays, vehicle attitude,and robot positioning, meets the sun protection requirements. Inaddition, the translation paths taken by the RS 106 when movinginstruments preferably will be bounded based on both sun protection andthermal constraints.

According to one embodiment, prior to using the DR 120 for servicing,the GA 108 retrieves the DR 120 from the EM 104. The remote team theninitiates a series of procedures to verify the performance of the DR 120and the total RS 106 system. Preferably, the GA 108, and the DR 120arms, 134 and 136, are moved sequentially. In one embodiment, the RS 106control system does not allow simultaneous motion of any arms 108, 134or 136. However, one skilled in the art will see that such simultaneousmotion, and other types of synchronized motion, are contemplated by thisinvention and fall within its scope. Thus, according to one preferredembodiment, in order to move one of the arms GA 108, 134 or 136, motorpower is disabled for the other two. In this embodiment, removing motorpower engages the brakes for that motor. Also, the command for puttingon the brakes can be part of a script, or can be a single command issuedfrom the remote operator, or can be part of an electrical orelectromechanical interlock that automatically engages to preventsimultaneous motion of two or more of the arms GA 108, 134 or 136.

As indicated below, two tests preferably are carried out on the DR 120to verify readiness, include the following steps.

DR Aliveness Test

-   -   Once the GA 108 and the DR 120 are mated, heater power is        applied to the DR 120 until it reaches operating temperature.        This can be verified via EM 104 bay temperature telemetry.    -   The DR 120 VCUs are powered and the flight software is loaded        into the DR 120 flight computers.    -   Communication between the DR 120 computers and the VCUs is        verified.    -   All DR 120 joints are verified by issuing commands for small,        benign movements to each joint against the DR 120 launch locks.    -   The DR 120 down-locks are released and the DR 120 is retracted        by the GA 108 from the EM 104 and moved to a hover position away        from the structure.        DR Functional Test    -   A VCU test verifies the use of each DR 120 camera in still and        streaming mode, as applicable. All DR 120 cameras 130 and 132        are tested in both still and streaming mode. In addition, all DR        120 light sources are tested, which ensure proper video and        still photo capabilities for both teleoperation and supervised        autonomy.    -   The joint range of motion is verified by commanding each joint        through the widest practical range and polarity.    -   The dexterous motion of the DR 120 is verified by moving to        several preplanned locations that are away from any hardware to        test the inverse kinematics and singularity avoidance software.        Verification is accomplished via joint angle telemetry and        streaming video feeds.    -   Two performance tests are executed to verify the transient        response of the DR 120 arms 134 and 136 in several poses and        modes. Various disturbances are introduced and arm performance        is verified via camera views and joint angle sensors. The first        test, summarized in Table 10 below, is the DR transient response        test. Each test category is repeated with the opposite DR 120        arm stabilized and free. The second performance test, summarized        in Table 11 below, verifies GA 108 transient response while        mated to the DR 120.    -   The OTCM 138 cameras are calibrated using a calibration fixture        on the task board. The DR 120 arms 134 and 136 then translate to        the high hover position directly over the calibration fixture.        Still images are downlinked and compared to expected image        fidelity.

TABLE 10 Test Disturbance Category Arm Pose DR Arm Mode Input Objective1 Various elbow joint Position Hold Fire Thruster Damping, 2 anglesBrakes Engaged Fire Thruster natural frequency, 3 (outstretched, rightPosition hold Arm slew link flexibility, angle, folded up) (in freedrift) joint flexibility. 4 Stretched Out Brakes Engaged Fast arm slewStopping (in free drift) Distance 5 Various approaches Tele-operated.Small arm slew Positioning to Task Board. (in free drift) resolution,ergonomics

TABLE 11 Test GA Arm Disturbance Category GA Arm Pose Mode InputObjective 1 Two nominal servicing Position Hold Fire Thruster Damping, 2configurations: One Brakes Engaged Fire Thruster natural frequency, 3 atEM, other at HST Position hold Arm slew link flexibility, (in freedrift) joint flexibility. 4 Stretched Out Brakes Engaged Fast arm slewStopping (in free drift) Distance 5 Various approaches to Tele-operated.Small arm slew Positioning HST Bay. (in free drift) resolution,ergonomics

Once the OTCM 138 cameras are calibrated, a test of the ObjectRecognition and Pose Estimation (ORPE) system can be done. Preferably, aDR 120 arm 134 or 136 translates to a hover position directly over thetask board test area. A stereo pair of images is taken by the calibratedand synchronized OTCM 138 cameras, and these images are then downlinked,and the OTCM offset from the calibration fixture is calculated. Acommand is then sent to correct the arm position for the offset. Oncecamera calibration is complete, the DR 120 is commanded to move by apredetermined amount, and the ORPE process is repeated to ensure ORPEfunctionality. Other tests that optionally may be done at this timeinclude a force moment accommodation test (to compensate for forces andmoments placed on the DR 120 when transporting tools and instrumentsduring a task), an OTCM torquer test (to verify the running torqueprofile of the OTCM torque drive), and an OTCM umbilical connectorcheckout (to power payloads in its grasp via an umbilical connector)

The preferred RS 106 motion might be categorized into two distincttypes: constrained movement and free-space movement. Constrained motionis defined as any task that occurs in close proximity to otherstructures, including the EM 104, DM 102 and HST 50. Free-space motionis movement that takes place away from a structure, such as the movementof the RS 106 from one worksite location to another, or the movement ofthe DR 120 to a grapple fixture hover position. Preferably, thedelineation between constrained and free-space motion is based on themaximum braking distances of the GA 108 and DR120.

Free-space motion is preferably done using supervised autonomy(described in more detail below). In this embodiment, GA 108translations between the EM 104 and the HST 50 preferably are consideredfree-space motions. Other examples might include visual surveys andcoarse positioning. In one embodiment, visual surveys will be requiredthe first time the DR 120 visits a new site or when the DR 120 returnsto the worksite after a significant absence. These return visual surveysare intended to be a quick visual check of the area with the purpose ofverifying that there has been no change to the worksite. The visualsurveys in general may be accomplished using any of the robot cameras,depending on the size of the area to be surveyed and lightingconditions.

In another embodiment, coarse positioning is described as a series ofmaneuvers to position the RS 108 for a task. The goal is to position oneDR 120 end effecter approximately 30 cm from the intended interface,such as a micro-fixture, while the other DR 120 end effecter ispositioned to obtain the necessary orthogonal camera views. FIG. 24illustrates typical arm positioning after completion of coarsepositioning. Preferably, this pose, at a distance of about 30 cm, iscalled a “high hover” position.

Constrained motion, according to one embodiment, is done using acombination of supervised autonomy and teleoperation, depending on thetask interface (described hereafter). Fine positioning, contactoperations and Object Recognition/Pose Estimation (ORPE) are allexamples of constrained motion. Once the arms are sited using coarsepositioning, ORPE can be used to account for any misalignment that mayexist between the end effector and the micro-fixture before contact isattempted. Preferably, arm motion is halted until this alignment iscomputed.

As one skilled in the art will realize, the fluidity of operations inmanual mode is dependent on the latency of the system, which can bedefined as the time from when the command to move is issued from theremote operator via a hand controller until the operator receivesverification of the motion via video. The RV 100 on orbit and remotesegments preferably are designed to minimize this latency and also tominimize variations in latency. According to one embodiment, thistransmission latency is about two seconds.

Supervised autonomy as it applies to the RS 106, is defined as theprocess of allowing the RS 106 to execute a sequence of instructionswithout intervention from the remote operator. Preferably, theseinstructions are packaged in a script that is generated remotely, and isuplinked to the RS 106 for execution upon ground command givingauthority to proceed. Alternately, RS 106 may execute a series ofoperations based on AI logic that may be preprogrammed into RS 106 priorto launch or uplinks or updated once RS 106 in orbit. The remoteoperator maintains the ability to abort a script or AI initiatedoperation at any time.

In one embodiment, two remote operators work in tandem as anoperator/co-operator team to “teleoperate” (manually control from afar)the RS 106. Preferably, during tasks using supervised autonomy, bothoperators will monitor script/AI execution and resulting RS 106 motion.During teleoperation, or manual operation, the RS 106 primary operatorpreferably controls the GA 108 and both DR 106 arms 134 and 136 insequence as required for a servicing task. The other operator preferablysupports the primary RS 106 controller by navigating through theservicing procedures, helping to coordinate camera views, inspecting theworksite during a task, and so forth. One of skill in the art willrecognize that the above division of labor between two operators andthat a different division of tasks, completion of all tasks by oneoperator, or the use of more than two operators are all contemplated bythe present invention and fall within its scope. Tele-operation isexplained in greater detail hereafter.

It is preferred that supervised autonomy will be used for free-spacemotions. It is also preferably used for constrained motions when thetarget interface is robot-friendly, that is, when it is equipped with afixture and a target designed for robot operations. Examples ofrobot-friendly operation include removing/returning tools from/tocaddies, handling RV 100 hardware, and interfacing with HST 50 when tooldesign allows for self alignment onto the HST 50 interface without theuse of robot-friendly targets, such as clamping a tool onto an HSThandrail. A preferred decision tree for determining which tasks are notrobot-friendly, and therefore done by teleoperation, is provided in FIG.25.

Supervised autonomy, according to one embodiment, relies on the abilityto correlate RS 106 motion with an environment model at the remotestation. This is done through the use of Object Recognition/PoseEstimation (ORPE), which is described as the process of detecting errorsin and correcting the remote model upon which scripted commands arebased. Preferably, after the RS 106 reaches the high hover position overa target area (about 30 cm), three sets of stereo images are taken withthe cameras. These images are then downlinked to the remote station,preferably uncompressed (10 Mb for each set). Once the images are at theremote station, a team of image analysts studies the current positionand orientation with respect to the target, and compares this with theexpected model in the remote system software. This visual comparisonshows any misalignment between the arm's coordinate frame and the targetfixture coordinate frame.

According to one embodiment, if a misalignment is detected, a poseestimate is applied to the target coordinate frame via a transformationmatrix. The new pose estimation for the robot is transmitted to thevirtual environment software, which updates the target coordinate framewithin the model. This updated model is then placed on the server at theremote station and a notification is sent to all remote workstationsthat a model correction has been made. Personnel at each workstationmust acknowledge the change. Then a correction script is generated basedon the new model. Before the script can be uplinked, it must bevalidated against the new model version number. Any attempt to validatea script with an old model will be rejected by the remote systemsoftware, thus allowing for an extra check of the model. Once the scriptis validated, it is loaded to the spacecraft and executed. FIGS. 26A,26B and 26C illustrate the alignment before and after the correction isapplied. One of skill in the art will recognize that it also is possibleto perform that above model correction using processing on board RS 106.In one embodiment, the environmental model is stored on the RS 106 andthe correlation of RS 106 motion with that model may be automaticallyinitiated by an on board processor based on a command from the remotestation, a preloaded script, or AI logic resident in the on boardprocessor.

The arm then preferably moves to a new hover position in alignment withthe target and the task continues. In accordance with this embodiment,if the position and/or orientation errors detected by ORPE areunexpectedly large, a script simulation, or rehearsal, may be requiredat the remote station prior to uplink. If the errors are determined onboard RS 106, the on board processor can run a simulated rehearsal ornotify the ground operator of the need to perform a script simulation atthe remote station.

One preferred ORPE process can take anywhere from 10 to 15 minutes,depending on image quality and the amount of correction needed. Oneskilled in the art will understand that lighting is important for ORPEto function properly. Thus, the preferred lighting system used toilluminate the worksites is optimized to provide sharp contrasts thatcreate the ideal images for used by ORPE software.

As indicated earlier, a preferred RS 106 can also be operated via manualcontrol, also called teleoperation. An operator at a remote locationmonitors the environment via video downlink and moves the RS 106preferably using joystick control. Deflection of the hand controllersdirectly results in RS 106 movement.

This mode of operational control can be used as an alternative tosupervised autonomy for a given task and is the preferred primary methodfor positioning the DR 106 when the interface being engaged is notdesigned for robot use, as, for example, in the case of installation ofhandling tools on the science instruments for removal from HST 50. Inone embodiment, two hand controllers located on the remote controlstation are used to move the arms via rate commands, meaning that theamount of controller deflection equates to how fast the arm moves.Moreover, these commands can be scalable so the same amount ofdeflection of the controllers can generate large motion if the arms arefar from structure, or small motion if the task being carried outrequires very fine positioning. During teleoperations, the operatorpreferably relies primarily on streaming video from a variety ofcameras, including the arm's cameras, to detect misalignments andposition the DR 106 or a tool in the correct location.

Therefore, teleoperator commands are preferably verified visually.Whenever possible or practical the arm 134 and 136 cameras will offer anorthogonal view of the interface to be grasped, while cameras located onthe DR 106 body provide situational awareness views of the worksite.Preferably, camera switching will be done via a voice-loop request fromthe teleoperators at the remote station to camera operators at aseparate vision system workstation. In a preferred embodiment, theswitch between scripted and manual modes can be carried out quickly andeasily by commanding a mode change to the RS 106 avionics andconfiguring the remote station for commanding via the hand controllers.

One skilled in the art will appreciate that the servicing tasks arepreferably choreographed such that they can be carried out using onlyone DR 106 arm 134 or 136, thereby allowing the other arm to serve as astabilization point at a worksite (using appropriate stabilizationinterfaces). One embodiment makes use of this feature for servicingtasks where the dynamics of the system cause 1) inadequate positioningresolution with respect to task alignment requirements, 2) excessiveovershoot during positioning or stored energy release with respect toaccess envelopes, 3) excessive settling time with respect to allocatedtask durations, or 4) excessive system deflection when force is appliedwith respect to task alignment requirements. The use of stabilization inthis embodiment may decrease task completion times in some cases.

For each servicing task, a number of tools may be required to interfacewith HST 50, or other satellite serviced, to overcome the fact that thesatellite may not be equipped with robot-friendly interfaces. Forexample, in one embodiment, tools are used to open HST's aft shroud andradial bay doors, de-mate and mate connectors, and drive instrumentlatches. These tools preferably are stowed on the EM 104 in stowagecompartments, although it will be recognized that tools could also bestowed on the DM 102. Tools needed for each servicing task arepreferably grouped together and placed onto a caddy prior to launch. Forexample, the tools needed to open the aft shroud doors are all locatedon one caddy, and the tools needed to carry out the servicing tasks oninstruments are in another caddy. One preferred tool caddy includes aplate upon which the tools are attached with release mechanisms andtargets that allow for easy removal and replacement of tools by the DR120. These caddies save time because the RS 106 does not need to travelback and forth to the EM 104 to retrieve tools. Instead, as part of thesetup for a servicing task, the RS 106 removes the entire caddy from itsstowage location on the EM 104, transports the caddy to HST 50, andmounts it onto the worksite preferably using a Foot Restraint (FR)socket. Tools can be used multiple times during servicing and hence eachtool can be easily located by the DR 120 cameras through the use ofvisual targets and released from the caddy or replaced by incorporatingmicro-fixtures into their design.

In one preferred embodiment, certain servicing activities can continuethrough a loss of communications (command and telemetry) between theremote station and the RV 100. For example, during a TDRS Zone ofExclusion (ZOE), where no TDRS satellite is in sight, operations can beplanned accordingly. Thus, approaching a ZOE, a scripted command can besent with the resulting task continuing through the ZOE. Alternately, AIlogic on board RV 100 or any of its components can continue to initiatetasks autonomously through the ZOE. Preferably, the AI logic can bepreprogrammed with a plurality of safety criteria that must be metbefore initiating particular autonomous tasks. In one embodiment, oneset of safety criteria can be established to govern autonomous actionwhen communications are available, i.e. supervised autonomy, and asecond, preferably more stringent, set of safety criteria may beestablished for autonomous actions when communications are notavailable, i.e. unsupervised autonomy.

In the event of a temporary loss of signal (LOS), teleoperated armmovements will stop immediately with a loss of commanding. Supervisedautonomous movements on the other hand have the capability to bescripted to either continue through the LOS, as with a ZOE, or stopimmediately, depending on the task. Scripted activities are preferablyevaluated on a case by case basis to determine the level of riskassociated with the completion of the activities while there are noremote communications. In either scenario, the robot reports to theremote station upon retrieval of signal.

In one preferred embodiment, after the servicing is concluded and beforethe EM 104 separates from the DM 102, the serviced satellite may beboosted to a higher orbit if its orbit has deteriorated significantly.Reboost, in this case, would preferably be performed using the EM 104propulsion system. In another embodiment, the satellite being servicedmay be refueled by the EM 104 prior to separation to further extend thesatellite's useful life.

The EM 104 ejection and disposal phase begins, according to oneembodiment, when control authority is handed over from the DM 102 to theEM 104 and ends when controlled disposal repositioning of the EM 104 iscomplete. Alternately, the EM 104 may be separated from DM 102 andeither (1) directed to a second satellite requiringde-orbiting/servicing; or (2) placed in a parking orbit to awaitinstructions to proceed to a second satellite, optionally waitingserving as an on orbit repair facility to enable rapid return to serviceupon unexpected failure of a satellite or key component thereon.

FIG. 27 summarizes one embodiment of the EM 104 ejection and disposalphase. One embodiment of the sequence is as follows:

-   -   1. EM 104 ejection from the RV 100: After the RV 100 is slewed        to the proper attitude, release pyrotechnics are fired to        liberate the clamp band between the DM 102 and EM 104.        Preferably, springs create a relative separation velocity        between the EM 104 and the DM 102/HST 50 stack sufficient to        ensure that there is no re-contact during the time required for        initial EM 104 attitude determination and stabilization. One        embodiment releases the EM 104 directed along the negative orbit        velocity vector. Post-ejection, the EM 104 nulls its own tip-off        rates while the HST 50 nulls the tip-off rates of the DM 102/HST        50 stack. The EM 104 then performs two zenith-directed burns of        the thrusters followed by a negative orbit velocity vector burn.        Other separation mechanisms and methods known to those of skill        in the art may be used without departing from the scope of the        present invention. If EM 104 is to be placed in a parking orbit        or transit directly to a second satellite in need of repair, EM        104 may be boosted into that parking orbit or an intercept        trajectory to the second satellite following separation from the        DM 102/HST 50 stack.    -   2. EM 104 mass properties measurement: When the EM 104 is safely        in a non-interfering orbit with respect to HST 50, the EM 104        preferably will perform a series of RCS thruster firings in        order to determine its mass, center of mass, and moments of        inertia, and update the appropriate table. Remote station        control may verify proper operation. If EM 104 is to be        immediately de-orbited, steps 3 and 4 below may be executed. If        EM 104 is to be placed in a parking orbit for later servicing        mission(s) for one or more additional satellites, steps 3 and 4        may be performed at the conclusion the last servicing rendezvous        for EM 104.    -   3. De-orbit burn #1: At the appropriate point in the orbit to        place perigee (accounting for perturbations) for the final        disposal burn, the EM 104 will fire its four 100-pound thrusters        for sufficient time to produce a predetermined delta V, dropping        perigee to 250 km. This perigee is high enough to avoid attitude        control issues associated with center-of-mass/center-of-pressure        offsets. It is also high enough to allow the EM 104 to remain in        orbit long enough to recover from any anomalous condition.    -   4. De-orbit burn #2: At least 3 orbits later (after sufficient        time to confirm a nominal burn #1), the EM 104 will fire its        four 100-pound thrusters for sufficient time to result in        perigee below 50 km. There will be two consecutive orbits during        which burn #2 can be executed, with additional pairs of        opportunities each day.

The entire sequence described above is preferably autonomous, with theremote operator continuously monitoring performance and providinggo-aheads before burns #1 and #2. Any anomalies detected on boardpreferably result in an abort of the sequence, except during de-orbitburn #2, where a point of no return is reached. The remote operator canabort the sequence at any time prior to de-orbit burn #2 via a commandor EM 104 can be set automatically to abort the sequence if an anomalyis detected.

Preferably, after the EM 104 is clear of the HST 50/DM 102 stack, the DM102 solar arrays 142 are deployed. FIG. 28 graphically depicts thisprocess. The preferred science operations phase begins at the completionof EM 104 ejection and terminates when science observations and anyend-of-life testing is complete. According one embodiment, after the EM104 has been jettisoned the HST 50 is commanded from the remote stationto reconfigure for science operations. This includes establishingattitude control, deploying the HGA booms 56, opening the aperture door58, and so forth. One of skill in the art will recognize that, followingservicing, the satellite will undergo a plurality of processes to resumeits normal post-service operations, which may include establishing orrestoring attitude control, communications links, C&DH, thermal control,and payload operations on the satellite. The particular sequence of postservicing operations will vary depending on the configuration andpayload of the satellite as well as its condition when servicing wascommenced.

In another embodiment, the HST science operations will resume with theinitiation of the Servicing Mission Verification Program (SMOV). TheSMOV for the preferred servicing mission has been established to verifythe functions of the HST 50 replacement instruments. It also includesthe re-commissioning of the existing science instruments, spacecraftsubsystems, and the overall observatory for science operations. FIG. 29summarizes the science operations phase.

Due to the extended timeframe of the method of one embodiment, theon-orbit activities of the SMOV program can be grouped for execution intwo phases, thereby mitigating some of the HST observing time lost dueto a lengthy servicing mission. Preferably, activities in SMOV Phase Aare those that can be carried out while the EM 104 is still attached tothe HST 50. Activities executed during Phase A will complement theindividual hardware elements performance verification (FT of battery andgyro augmentations, and instruments). These may include, but not belimited to, engineering activation of old and new science instruments,monitoring their contamination and thermal properties, andcharacterizing their baseline performance. Wherever possible, sciencecalibrations (internal and external) are performed. Also, if at allpossible, science programs can be carried out during servicing, subjectto the constraints imposed by the servicing mission and pending thecommissioning of prerequisite capabilities. All activities performedduring Phase A will not be sensitive to and will not interfere with thetemporary spacecraft configuration and the established servicing missiontimeline.

SMOV Phase B preferably comprises those commissioning activities thatcan be carried out only after completion of the servicing mission andthe release of the EM 104, with the spacecraft in final on-orbitconfiguration. Thus, in one embodiment, commissioning of all otherscience instruments and spacecraft subsystems not performed during PhaseA will be accomplished in Phase B. This may include spacecraft power,pointing, thermal, and guidance, as well as existing and replacementscientific instrument characterization and calibration. Hence, scienceobservations preferably ramp up to their normal levels as SMOV Phase Bactivities ramp down to completion. The HST 50/DM 102 disposal phasepreferably begins at the completion of science operations and terminateswith the completion of the controlled disposal. FIG. 30 summarizes thisphase.

Preferably the DM 102 slews the combined vehicle to the preferredattitude for the initial reentry burn. According to this embodiment, asmall retrograde burn (engineering burn) to check out the system isinitiated followed by two retrograde apogee burns. Then the final burnadjusts the perigee to 50 km and sets up the preferred controlledreentry into the Pacific Ocean.

While the invention herein revealed and described is set forth in what,at present, is considered to be the best mode contemplated for makingand carrying out the invention and the preferred embodiments of thisinvention, it will be understood that the foregoing is given by way ofillustration, rather than by way of limitation. Accordingly, any and allboundaries and restrictions imputed to the scope of this invention mustbe defined by the spirit and intent of the following claims.

1. A servicing vehicle for servicing other free-flying spacecraft, comprising: a de-orbit module, and an ejection module including a robot system for servicing the spacecraft, the robot system having a grappling arm, and at least one dexterous arm, wherein the robot system is autonomously controlled by commands that are subject to override by a remote operator in telecommunication with the servicing vehicle.
 2. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the ejection module further comprises stowage area for storing the robot system.
 3. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the de-orbit module includes stowage area for tools, parts and instruments needed in servicing the spacecraft.
 4. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the de-orbit module includes: relative and absolute navigation sensors for determining absolute servicing vehicle attitude and relative attitude between the servicing vehicle and the other spacecraft, guidance, navigation and control system actuators, thrusters, and momentum management devices.
 5. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the ejection module further includes stowage area for tools, parts and instruments needed in servicing the spacecraft.
 6. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the grappling arm is used to capture the other free-flying spacecraft.
 7. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the de-orbit module and the ejection module are separable.
 8. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 1, wherein the robot system is used to carry out servicing tasks.
 9. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 8, wherein the servicing tasks include replacing degraded batteries with new batteries.
 10. A servicing vehicle for servicing other free-flying spacecraft, according to the limitations of claim 8, wherein the servicing tasks include replacing existing scientific instruments with new scientific instruments. 